In an aircraft gas turbine engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against the airfoil section of the turbine blades, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward. The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. Thus, there is incentive to raise the combustion gas temperature.
In the compressor portion of an aircraft gas turbine engine, atmospheric air is compressed to 10-25 times atmospheric pressure, and adiabatically heated to 800°-1250° F. in the process. This heated and compressed air is directed into a combustor, where it is mixed with fuel. The fuel is ignited, and the combustion process heats the gases to very high temperatures, in excess of 3000° F. These hot gases pass through the turbine, where rotating turbine wheels extract energy to drive the fan and compressor of the engine, and the exhaust system, where the gases supply thrust to propel the aircraft. To improve the efficiency of operation of the aircraft engine, combustion temperatures have been raised. Of course, as the combustion temperature is raised, steps must be taken to prevent degradation of engine components directly and indirectly as a result of the higher operating temperatures.
The requirements for enhanced performance continue to increase for newer engines and modifications of proven designs, as higher thrusts and better fuel economy are among the performance demands. To improve the performance of engines, the combustion temperatures have been raised to very high temperatures. This can result in higher thrusts and/or better fuel economy. These combustion temperatures have become sufficiently high that even superalloy components not within the combustion path have been subject to degradation. These superalloy components have been subject to degradation by mechanisms not previously generally experienced, creating previously undisclosed problems that must be solved. One recent problem that has been discovered during refurbishment of high performance aircraft engines has been the corrosive pitting of turbine disks, seals and other components that are supplied with cooling air. The cooling air includes ingested particulates such as dirt, volcanic ash, fly ash, concrete dust, sand and sea salt, as well as metal, sulfates, sulfites, chlorides, carbonates, various and sundry oxides and/or various salts in either particulate or gaseous form. These materials are deposited on substrate surfaces. When deposited on metallic surfaces, these materials can interact with one another and with the metallic surface to corrode the surface, which is accelerated at elevated temperatures. The materials used in turbine engines are typically selected on high temperature properties, including their ability to resist corrosion. Even these materials will degrade under severe conditions at elevated temperatures. On investigation of the observed pitting problem, it has been discovered that the pitting is caused by a formation of a corrosion product as a result of the ambient airborne foreign particulate and gaseous matter that is deposited on the disks, seals or other components as a result of the flow of cooling air containing it. This deposition, along with the more elevated temperature regimes experienced by these engine components, has resulted in the formation of the corrosion products. It should be noted that the corrosion products are not the result of exposure of the engine components to the hot gases of combustion, normally associated with oxidation and corrosion products from contaminants in the fuel. The seals, turbine disks and other components under consideration and discussed herein generally are designed so that if a leak is present, the air will leak in the direction of the flow of the hot gases of combustion and not in the direction of the components under consideration.
Because the corrosion products are the result of exposure of the engine components to cooling air drawn from ambient air environments, it is not uniform from engine to engine as aircraft visit different geographic locations with different and distinct atmospheric conditions. For example, some planes are exposed to salt water environments, while others may be subject to air pollutants from highly industrial regions. A variety of coatings have been developed to mitigate corrosion concerns.
Known coating systems, in addition to the coating for mitigating corrosion, hereinafter referred to as a “base coating,” also has included an inorganic top coating, such as a phosphate coating. These known systems suffer from the drawback that the top coat remains present at operating temperatures and interferes with the base coating. Specifically, when large temperature fluctuations occur, such as those seen in the engine components exposed to cooling air, the top coat results in degradation of the base coat, including potential failure within the coating, which increases the possibility of corrosion of the engine component. When the inorganic top coat, including phosphate sealant, is directly applied to the base coat, the sealant is believed to infiltrate the pores of the base coating reducing the ability of the base coat to tolerate strain, such as the strain experienced during thermal cycling (i.e., thermal cycles in excess of 500° F./min). In addition to the above, known top coats are difficult to apply, requiring masking and making of the material to produce the coating. For example, a known coating may require 10-12 hours or more to provide the inorganic top coat, wherein the delay adds to the total time and cost required for manufacture of the gas turbine engine.
Elimination of the top coating entirely is also undesirable. The use of a base coat alone (i.e., with no top coat), has a number of disadvantages. First, the base coat is a generally porous coating structure that is susceptible to discoloration and/or staining due to oils or greases that may contact the surface of the engine component during manufacture. Oil, grease or similar contaminants are drawn into the porous structure of the base coat and produce an undesirable appearance, due to the resulting discoloration and/or staining. In addition, the incorporation of contaminants into the porous structure may also reduce the ability to withstand temperature fluctuation and the corrosion resistance of the base coating. Second, the surface appearance of the engine component is generally a dull, matte texture and not aesthetically pleasing. Therefore, it is undesirable to provide components having only a base coating without additional top coatings.
What is needed is a coating system that provides a manufactured component having an aesthetically pleasing surface finish that is protected against surface damage during the manufacturing process and does not detrimentally affect the underlying coating when the turbine engine components are subjected to elevated operating temperatures and extreme changes in temperature in a wide variety of atmospheres. In addition, what is needed is a coating system that can be applied quickly and inexpensively, without the drawbacks of the inorganic top coats known in the art.